Turboprop-powered aircraft

ABSTRACT

A turboprop-powered medium altitude long endurance aircraft, having a gas turbine engine; a heat scavenging device to scavenge heat from the gas turbine engine; and a heating device to use the scavenged heat to provide heating to the aircraft. The heat scavenging device may be placed on an engine casing and/or on or in an engine exhaust duct. The heating device may include a circulation path routed directly to a location in the aircraft where heating is to be performed, for example a leading edge of an engine support pylon or a leading edge of an engine-carrying wing. The heating device can include a heat exchanger.

FIELD OF THE INVENTION

The present invention relates to thermal management of a turbopropaircraft, in particular a turboprop-powered Medium Altitude LongEndurance (MALE) aircraft.

BACKGROUND

For aircraft in general, many types of thermal management arrangementsare known.

On the one hand, many types of aircraft and/or aircraft mission profilesare relatively simple, for example light aircraft with piston-poweredpropellers flying at low speeds at low altitude, and the thermalmanagement arrangements may accordingly be relatively undemanding, e.g.use of simple ram air may suffice.

On the other hand, many other types of aircraft and/or aircraft missionprofiles have highly demanding thermal management requirements, butfactors such as the cost and complexity of those aircraft, and/or thehigh availability of fuel driven power during flight, make providingcorrespondingly complex and expensive (in terms of money or powerconsumption) thermal management arrangements practicable. For example,in jet-powered aircraft it is practicable to provide significantelectrical power for heating and it is also practicable to use bleedpower from the jet engines to provide heating or (when cooling isrequired) power to drive a conventional air conditioning system.

One specific type of aircraft with an associated mission profile is aturboprop-powered Medium Altitude Long Endurance (MALE) aircraft.Typically, a turboprop-powered MALE aircraft may carry out flights oflong duration, e.g. more than 12 hours or even more than 24 hours oreven more than 36 hours, and for example may fly at relatively highaltitudes of between 30,000 and 50,000 feet (i.e. high compared to otherknown light aircraft) where there are lower temperatures, and may be arelatively cheap or simple aircraft. Also, a turboprop-powered MALEaircraft may be required to operate in areas of high temperature atground level (e.g. up to 55° C.) and at low altitudes during take-offand approach.

SUMMARY OF THE INVENTION

The present inventors have realised that known aircraft thermalmanagement arrangements are not ideally suited for a turboprop-poweredMALE aircraft, in view of the above described characteristics of aturboprop-powered MALE aircraft and its mission profile, when consideredindividually and/or in combination. In particular, the present inventorshave realised that the thermal management requirements for aturboprop-powered MALE aircraft fall in a middle-ground that is notsuitably addressed by either of the extremes (undemanding or highdemanding) of known thermal management arrangements mentioned earlierabove. Consequently the present inventors have realised that thefollowing may be beneficial either individually or in combination.

The present inventors have also realised that in view of the longendurance requirement for a turboprop-powered MALE aircraft it would bedesirable to provide a means for capturing heat (to e.g. use to heat thepayload when flying at altitudes that are cold) that reduces or removescompletely any need to extract heat from the power being provided forpropulsion (which would otherwise be the case conventionally with e.g.taking bleed air from the gas turbine engine or with e.g. extractingshaft power). Thereby the power would be preserved for contributing toextended flight endurance.

The present inventors have also realised it would be desirable toprovide a cooling arrangement that does not need to use active airconditioning.

The present inventors have also realised it would be desirable toprovide an intermediate cooling process for an intermediate part of themission profile. In more detail, the present inventors have realisedthat in hot locations, the mission profile may be considered, forcooling purposes, as comprising various stages, for example a firststage where the aircraft is stationary on the ground where aconventional ground cooling trolley can be connected to the aircraft andcarry out cooling of e.g. the aircraft payload, a second stage startingfrom when where the aircraft is disconnected from the cooling trolleyand ending when the aircraft has climbed to a height where the ambientair is sufficient to be used to perform the required cooling, and athird stage when the aircraft is flying at or above the height requiredfor the ambient air to be sufficient to be used to perform the requiredcooling. Note during the second stage the aircraft may be required to bestationary, perform taxiing to a take-off position, perform take-off andthen climb to the later-mentioned altitude. Using this terminology, thepresent inventors have realised that cooling during the first and thirdstages can be performed relatively simply, whereas cooling during thesecond stage is more challenging. Accordingly, the present inventorshave realised it would be desirable to provide an intermediate or“stop-gap” cooling process for the second stage that does not undulycomplicate the simple cooling processes that may be used for the firstand third stages. The present inventors have further realised that, dueto the intermediate or “stop-gap” nature of such a cooling process forthe second stage, it would be desirable to provide a solution for thesecond stage that is particularly efficient during its limited time ofdeployment, and where that particular efficiency was derived at least inpart by making use of the very fact that it is only required to carryout its cooling role for a short limited duration of the second stage,and that moreover could preferably be easily “reset” during e.g. thefirst stage of the turboprop-powered MALE aircraft's next flight and/orduring the third stage of a flight. The present inventors have furtherrealised that such a solution could be provided by employing materialsof e.g. high specific heat capacity/latent heat selected to have desiredphase changing point temperatures, e.g. melting point. The presentinventors have further realised it would be desirable to provide novelheat exchanger apparatus or arrangements to provide particularlyefficient ways of using the materials mentioned in the precedingsentence, and that are particularly suitable for use on aturboprop-powered MALE aircraft.

The present inventors have further realised it would be desirable toprovide a thermal management arrangement for a turboprop-powered MALEaircraft that is able to provide various of the above solutions in aturboprop-powered MALE aircraft such that they can be selectedindividually or in combination according to particular operatingconditions a turboprop-powered MALE aircraft undergoes on the same ordifferent flights.

In a first aspect, the present invention provides a turboprop-poweredmedium altitude long endurance aircraft, comprising: a gas turbineengine arranged to provide propulsion power to an associated propeller;heat scavenging apparatus arranged to scavenge heat from the gas turbineengine; and heating apparatus arranged to use the scavenged heat toprovide heating to the aircraft.

The heat scavenging apparatus may comprise a thermal blanket.

The thermal blanket may comprise capillary tubing.

The heat scavenging apparatus may be placed on the engine casing of theengine.

The heat scavenging apparatus may be placed on or in an exhaust duct ofthe engine.

The heating apparatus may comprise a circulation path routed directly toa location in the aircraft where heating is to be performed.

The location in the aircraft where heating is to be performed may be aleading edge of an engine support pylon or a leading edge of anengine-carrying wing.

The heating apparatus may comprise a heat exchanger.

The heat exchanger may be arranged to use the scavenged heat to heat airin a recirculating air path.

The recirculating air path may be arranged to provide cooling and/orheating to the aircraft.

The turboprop-powered medium altitude long endurance aircraft mayfurther comprise a further heat exchanger; wherein: the further heatexchanger comprises a heat-storage material; the further heat exchangeris arranged to cool air recirculating around the air recirculation path;and the air recirculation path is arranged to allow the recirculatingair to provide cooling to the MALE aircraft.

The further heat exchanger may be arranged to allow the heat-storagematerial to be cooled by ground-based cooling apparatus when theaircraft is on the ground.

The further heat exchanger may be arranged to allow the heat-storagematerial to be cooled by ram air when the aircraft is in flight.

The heat-storage material may have a melting point selected so as to berendered solid during the cooling by the ground-based cooling apparatusand/or during the cooling by the ram air.

The heating apparatus may comprise a pressurised heat transfer liquidloop.

The pressurised heat transfer liquid loop may be arranged to transferscavenged heat to parts of the aircraft to perform icing prevention atthose parts of the aircraft.

The aircraft may be a surveillance aircraft and/or it may be an unmannedaerial vehicle.

In a further aspect, the present invention provides a method ofthermally managing a turboprop-powered medium altitude long enduranceaircraft, the method comprising performing heating of theturboprop-powered medium altitude long endurance aircraft using heatbeing scavenged from an engine of the aircraft.

The method may further comprise: using ground-based cooling air to coola heat exchanger, the heat exchanger comprising a heat-storage material;thereafter, using recirculated air passed via the heat exchanger to coolthe turboprop-powered medium altitude long endurance aircraft; andthereafter, using ram air to cool the heat exchanger.

Conventional, sophisticated aircraft, burdened with a significantpayload of electronic systems and sensors such as large scalesurveillance aircraft or electronic combat and communications systemsaircraft have merited associated sophisticated cooling/heating systems,with cabin pressurisation in the case of manned aircraft, and the enginepower-offtake for these systems has been tolerated.

It is desirable for such aircraft to be able to fly at higher altitudesat slower speeds, for example to undertake surveillance functions moreeffectively. Use of a turboprop-powered propulsion system for theaircraft optimises/maximises the fuel economy and hence the duration ofeach mission. However, the resulting combination of a large equipmentconditioning demand and a need to operate at low speeds at medium/highaltitudes can create unacceptable power/air offtake demands on turbopropengines both on the ground and in flight. In many cases, the aircraftmay be unmanned and so it is no longer necessary to provide theenvironmental conditions required when aircrew are present, makingsimpler system solutions possible.

By providing the aforementioned aspects of the present invention, thethermal environment required by the aircraft itself and/or theelectronic systems and sensors mounted in the equipment bay to operatereliably and effectively can be maintained without additional fuelconsumption or other power-offtake penalties in a very simple andflexible way. Any fuel payload can, therefore, be reserved for thepropulsion unit of the aircraft to extend the duration and/or complexityof the mission to be undertaken.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration (not to scale) of a turboprop-poweredMALE aircraft comprising a thermal management system;

FIG. 2 is a schematic block diagram of a thermal control system thatforms part of the thermal management system;

FIG. 3 is a schematic illustration (not to scale) showing certaindetails of a heat exchanger;

FIG. 4 shows a cross-section X-Y from FIG. 3;

FIG. 5 shows certain details of a water extractor that is part of theheat exchanger of FIG. 3;

FIG. 6 is a schematic illustration of a first example of a missionprofile that the turboprop-powered MALE aircraft may undergo;

FIG. 7 is a process flowchart showing certain steps performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft during the above mentioned first exampleof a mission profile;

FIG. 8 is a schematic illustration of a second example of a missionprofile that the turboprop-powered MALE aircraft may undergo;

FIG. 9 is a process flowchart showing a certain step performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft during the above mentioned secondexample of a mission profile;

FIG. 10 is a schematic illustration of a third example of a missionprofile that the turboprop-powered MALE aircraft may undergo; and

FIG. 11 is a process flowchart showing certain steps performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft during the above described third exampleof a mission profile.

DETAILED DESCRIPTION

FIG. 1 is a schematic illustration (not to scale) of a turboprop-poweredMALE aircraft 1 comprising a first embodiment of a thermal managementsystem.

The turboprop-powered MALE aircraft 1 comprises a fuselage 2, wings 4and engine support pylons 6.

Gas turbine engines 8, coupled to propellers 10, are attached to theengine support pylons 6. The gas turbine engines 8 drive the propellers10 to provide propulsion to make the turboprop-powered MALE aircraft 1travel in a flight direction 12. Each gas turbine engine 8 is encased inan engine casing 14, which itself is encased in a nacelle (not shown).Each gas turbine engine 8 has one or more exhaust ducts 16 (for clarity,in FIG. 1 only one exhaust duct 16 is shown for each gas turbine engine8).

The turboprop-powered MALE aircraft 1 further comprises an equipment bay18, which will typically carry a payload 20. The payload 20 may compriseone or more of surveillance equipment, navigation equipment,communications equipment, sensors, weapons systems, and so on.

The turboprop-powered MALE aircraft 1 further comprises a thermalcontrol system 22.

In this embodiment, the turboprop-powered MALE aircraft 1 furthercomprises an air inlet 24, an air outlet 25, an air circulation path 26,and a pressurised water-glycol heat transfer loop 28, each of which arecoupled to the thermal control system 22.

The air inlet 24 is arranged to input external air into the thermalcontrol system. In this embodiment, the air inlet 24 is arranged toprovide both ram air scoop functionality when in flight and an input forcool air from a ground cooling trolley (or other ground-based coolingapparatus) when on the ground (however, in other embodiments, twoseparate inlets may be provided, one as a ram air scoop, the other as ininlet for cool air from a ground cooling trolley). The air outlet 25 isarranged to output air after it has passed through the equipment bay. Aswill be explained in more detail later below, this is only used when itis not desired to recirculate the return air around the air circulationpath 26 (via the thermal control system 22).

The air circulation path 26 is arranged to input air from the thermalcontrol system 22 into the equipment bay 18 and to return the inputtedair out of the equipment bay back to the thermal control system 22. Inthis embodiment this is implemented by the air circulation path 26comprising a feed stage 30 and a return stage 32. In this embodiment thefeed stage 30 terminates in an open-ended aperture that releases inputair 34 (the term “input” being applied here to the equipment bay 18rather than the thermal control system 22) into the equipment bay 18,and the return stage has an open-ended return aperture that allowsoutput air 36 (the term “output” being applied here to the equipment bay18 rather than the thermal control system 22) to return to the thermalcontrol system. As will be explained in more detail later below, the aircirculation path 26 is operated for part of the turboprop-powered MALEaircraft's mission profile in a non-recirculating manner, and foranother part of the turboprop-powered MALE aircraft's mission profile ina recirculating manner.

The pressurised water-glycol heat transfer loop 28 is a closed looprecirculating liquid flow arrangement in which a water-glycol liquidmixture is kept pressurised. The pressurised water-glycol heat transferloop 28 comprises respective heat pick up stages 38 located within eachengine nacelle, a feed stage 40 arranged for the water-glycol to flowfrom the heat pick up stages 38 to the thermal control system 22, and areturn stage 42 arranged for the water-glycol to flow from the thermalcontrol system 22 to the heat pick up stages 38.

In this embodiment, thermal blankets 44 are fixed to one or moreselected areas of both the engine casings 14 and exhaust ducts 16. (Inother embodiments they may be fixed to one or more areas of only theengine casings 14, or to one or more areas of only the exhaust ducts, oradditionally or instead may be fixed to any other areas that will becomehot during operation of the turboprop-powered MALE aircraft such thatheat may be scavenged.)

In this embodiment the thermal blankets 44 comprise capillary tubingconnected to the pressurised water-glycol heat transfer loop 28 suchthat in operation the water-glycol mixture flowing through thepressurised water-glycol heat transfer loop 28 passes through thecapillary tubing (i.e. the thermal blankets are in effect part of thepressurised water-glycol heat transfer loop 28) and is heated by the hotengine casings 14 and/or exhaust ducts 16. The thermal blankets 44 alsocomprise insulation arranged over the capillary tubing to substantiallyretain within the pressurised water-glycol heat transfer loop 28 theheat gained captured by the capillary tubing.

The thermal blankets 44 are thermally coupled to the heat pick up stages38 of the pressurised water-glycol heat transfer loop 28. Thus, in thisembodiment, in operation, heat may be scavenged by the thermal blankets44 from the hot areas of the engine casings 14 and/or exhaust ducts 16,and transported via the heat pick up stages 38 and feed stage 40 to thethermal control system 22 which can then use or redistribute the thermalenergy as desired, and as will be explained in more detail later below.It is further noted that such flow of heated water-glycol from the heatpick up stages 38 to the thermal control system 22 can, if desired, also(or instead) be used to directly heat structures of theturboprop-powered MALE aircraft 1 that lie along the route of the feedstage 40 (or other provided routing), as will be discussed in moredetail later below.

The thermal management system of this embodiment comprises the thermalcontrol system 22, the air inlet 24, the air circulation path 26, thepressurised water-glycol heat transfer loop 28, and the thermal blankets44.

FIG. 2 is a schematic block diagram of the thermal control system 22.

In this embodiment, the thermal control system 22 comprises three heatexchangers, which for convenience will be termed the first heatexchanger 51, the second heat exchanger 52, and the third heat exchanger53.

The first heat exchanger 51 comprises an input 56 and an output 58 thatare operably coupled to (or may alternatively be considered as part of)the air circulation path 26.

The second heat exchanger 52 comprises a first input 60 and a firstoutput 62 that are arranged to form part of one or more further fluidcirculation paths, of which one such further fluid circulation path 80is shown in FIG. 2. The second heat exchanger 52 further comprises asecond input 64 and a second output 66 that are operably coupled to (ormay alternatively be considered as part of) the air circulation path 26.

The air circulation path comprises various valve systems which can beoperated to change the routes in which the air flows in the aircirculation path, as will be explained in more detail later below. FIG.2 shows five such valve systems, namely a first valve system 81, asecond valve system 82, a third valve system 83, a fourth valve system84, and a fifth valve system 85.

The first valve system 81 is positioned after the air inlet 24 at a3-way junction in the air circulation path 26. In operation, the firstvalve system 81 can be switched between either: (i) an off condition,which blocks air entering from the air inlet 24, or (ii) a first flowcondition in which air entering from the air inlet 24 is allowed to passalong a section of the air circulation path 26 to a 4-way junction inthe air circulation path 26 at which the second valve system 82 ispositioned, or (iii) a second flow condition in which air entering fromthe air inlet 24 is allowed to pass along a different section of the aircirculation path 26 to another 3-way junction in the air circulationpath 26 at which the third valve system 83 is positioned, or (iv) athird flow condition in which both the air flows of the first flowcondition and the second flow condition are allowed to take place at thesame time.

As mentioned above, the second valve system 82 is positioned at a 4-wayjunction in the air circulation path 26, where one of the sections ofthe air circulation path 26 meeting at the 4-way junction comes from thefirst valve system 81. The other sections meeting at the 4-way junctionof the second valve system 82 are a section coming from the output 58 ofthe first heat exchanger 51, a section coming from the second output 66of the second heat exchanger 52, and finally the feed stage 30 of theair circulation path 26 that feeds air into the equipment bay 18. Inoperation, the second valve system 82 can be switched between either:(i) a first flow condition in which air entering from the air inlet 24and then passing through the first valve system 81 to the second valvesystem 82 is allowed to pass into the feed stage 30 and hence into theequipment bay 18, or (ii) a second flow condition in which air flowingfrom the first heat exchanger 51 to the second valve system 82 isallowed to pass into the feed stage 30 and hence into the equipment bay18, or (iii) a third flow condition in which air output from the secondoutput 66 of the second heat exchanger 52 is allowed to pass into thefeed stage 30 and hence into the equipment bay 18.

As mentioned above, the third valve system 83 is positioned at a 3-wayjunction in the air circulation path 26, where one of the sections ofthe air circulation path 26 meeting at the 3-way junction comes from thefirst valve system 81. The other sections meeting at the 3-way junctionof the third valve system 83 are a section coming from a 3-way junctionwhere the fourth valve system 84 is positioned, and a section leadinginto the input 56 of the first heat exchanger 51. In operation, thethird valve system 83 can be switched between either: (i) a first flowcondition in which air entering from the air inlet 24 and then passingthrough the first valve system 81 to the third valve system 83 isallowed to pass into input 56 and hence into the first heat exchanger51, or (ii) a second flow condition in which air entering from the airinlet 24 and then passing through the first valve system 81 to the thirdvalve system 83 is allowed to pass on to the fourth valve system 84(from where it will pass into the second heat exchanger 52), or (iii) athird flow condition in which air coming from the fourth valve system 84is allowed to pass into input 56 and hence into the first heat exchanger51.

As mentioned above, the fourth valve system 84 is positioned at a 3-wayjunction in the air circulation path 26, where one of the sections ofthe air circulation path 26 meeting at the 3-way junction comes from thethird valve system 83. The other sections meeting at the 3-way junctionof the fourth valve system 84 are a section coming from a 3-way junctionwhere the fifth valve system 85 is positioned, and a section leadinginto the second input 64 of the second heat exchanger 52. In operation,the fourth valve system 84 can be switched between either: (i) a firstflow condition in which air entering from the air inlet 24 and thenpassing through the first valve system 81 and the third valve system 83to the fourth valve system 84 is allowed to pass into the second input64 and hence into the second heat exchanger 52, or (ii) a second flowcondition in which air coming from the fifth valve system 85 is allowedto pass on to the third valve system 83 (from where it will pass intothe first heat exchanger 51), or (iii) a third flow condition in whichair coming from the fifth valve system 85 is allowed to pass into thesecond input 64 and hence into the second heat exchanger 52.

As mentioned above, the fifth valve system 85 is positioned at a 3-wayjunction in the air circulation path 26, where one of the sections ofthe air circulation path 26 meeting at the 3-way junction goes to thefourth valve system 84. The other sections meeting at the 3-way junctionof the fifth valve system 85 are the return stage 32 of the aircirculation path 26 that returns air from the equipment bay 18, and asection leading to the air outlet 25. In operation, the fifth valvesystem 85 can be switched between either: (i) a first flow condition(which provides recirculation) in which air coming from the return stage32 is allowed to pass on to the fourth valve system 84 (from where itwill pass into either the second heat exchanger 52 or the first heatexchanger 51), or (ii) a second flow condition (when recirculation isnot taking place) in which air coming from the return stage 32 isallowed to pass on to the air outlet 25 and hence exits the aircirculation path 26.

In operation, the valve systems are switched as required between theirdifferent possibilities to provide one or more of the following modes ofair flow operation through the air circulation path 26:

(i) the first heat exchanger 51 may be used to cool air that is passedthrough it, to provide cooling air to the equipment bay 18, with the airbeing input continuously from the air inlet 24, with air that has passedonce through the equipment bay 18 being allowed to exit via the airoutlet 25, i.e. recirculation is not used;

(ii) the first heat exchanger 51 may be used to cool air that is passedthrough it, to provide cooling air to the equipment bay 18, with the airbeing recirculated around the air circulation path 26 (i.e. the fifthvalve system 85, the fourth valve system 84 and the third valve system83 are set to provide the condition in which air that has been passedthrough the equipment bay is returned to re-enter the first heatexchanger 51);

(iii) the second heat exchanger 52 may be used to heat air that ispassed through it, to provide heating air to the equipment bay 18, withthe air being input continuously from the air inlet 24, with air thathas passed once through the equipment bay 18 being allowed to exit viathe air outlet 25, i.e. recirculation is not used;

(iv) the second heat exchanger 52 may be used to heat air that is passedthrough it, to provide heating air to the equipment bay 18, with the airbeing recirculated around the air circulation path 26 (i.e. the fifthvalve system 85 and the fourth valve system 84 are set to provide thecondition in which air that has been passed through the equipment bay isreturned to re-enter the second heat exchanger 52).

The third heat exchanger 53 comprises a first input 68 and a firstoutput 70 that are operably coupled to (or may alternatively beconsidered as part of) the pressurised water-glycol heat transfer loop28. The third heat exchanger 53 further comprises a second input 72 anda second output 74 that are arranged to form part of one or more furtherfluid circulation paths. In this embodiment, one such furthercirculation path is the previously mentioned further fluid circulationpath 80 that also enters and exits the second heat exchanger 52. Thus,in this embodiment, when the valve systems 81-85 of the air circulationpath 26 are set so that the second heat exchanger 52 is used to provideheat to the air for heating the equipment bay 18, the heat is suppliedfrom the third heat exchanger 53, which has extracted such heat from thepressurised water-glycol heat transfer loop 28, i.e. such heat hasoriginally been scavenged by the thermal blankets 44 from hot parts ofthe engine casings 14 and/or exhaust ducts 16.

FIG. 3 is a schematic illustration (not to scale) showing certaindetails of the first heat exchanger 51. The first heat exchanger 51comprises a laminated stack of air flow channels 90 interlaced with heatsink layers 92. The first heat exchanger 51 further comprises arespective thermoelectric devices 94 positioned between the air flowchannels 90 and the heat sink layers 92. The laminated stack issurrounded by an insulation jacket 96. The input to the insulationjacket/laminated stack is provided by the earlier described input 56.The output of the insulation jacket/laminated stack feeds into a waterextractor 98. The output of the water extractor 98 is the earlierdescribed output 58.

In operation, input air 100 flows into the first heat exchanger 51, thenflows (as indicated by reference numeral 102 in FIG. 3) along the airflow channels 90, then exits the airflow channels 90 as cooled outputair 104. The cooled output air 104 then flows through water extractor 98and then exits as cooled and dried output air 106.

More details of the air flow channels 90 and heat sink layers 92 are asfollows. The air flow channels 90 are filled with open-cell aluminiumfoam. The heat sink layers 92 comprise open-cell aluminium foam combinedwith heat-storage wax, which is able to take in and store heatefficiently, e.g. without a large rise in temperature, due to highspecific heat capacity or to latent heat, selectable at a suitablemelting point (or put more generally, employing materials of e.g. highspecific heat capacity/latent heat selected to have desired phasechanging point temperatures, e.g. melting point). Any suitableheat-storage wax may be used, for example paraffin wax. The open-cellaluminium foam and the heat-storage wax are combined in any suitablemanner. The terminology “combined with” as used here should beunderstood to mean physically mixed, integrated, dispersed or thelike—i.e. the open-cell aluminium foam occupies, in a three-dimensionalmesh or cell-like formation, a certain amount of space within theoverall space defined by the heat sink shape, and the wax fills some orall of the remaining space, i.e. fills or partly occupies some or all ofthe voids within the aluminium foam structure.

A cross-section X-Y from FIG. 3 is shown in FIG. 4. More particularly,the cross-section of FIG. 4 shows certain details of the arrangement ateach interface between an air flow channel 90 and a heat sink layer 92.At the interface of the heat sink layer 92 and the air flow channel 90,the heat sink layer abuts one side of the thermoelectric device 94, andthe air flow channel 90 abuts the other side of the thermoelectricdevice 94. The thermoelectric device comprises a core part 110, and fins112 that extend from the core part 110 on either side of the core part110. The core part 110 is planar i.e. parallel to the planes of the heatsink layer 92 and the air flow channel 90. The fins 112 extend outwardsfrom the core part 110. Some of the fins 112 extend from one of theplanar surfaces of the core part 110 (the surface indicated by referencenumeral 114 in FIG. 4) into the heat sink layer 92. The other fins 112extend from the other of the planar surfaces of the core part 110 (thesurface indicated by reference numeral 116 in FIG. 4) into the air flowchannel 90.

In operation, during taxi/take-off/initial climb/approach/landing/taxi,heat is extracted from the air 102 flowing along the air flow channel 90and transferred via the thermoelectric device 94 into the heat sinklayer 92. Thus, in operation, the planar surface 114 operates as a coldface, and the planar surface 116 operates as a hot face. An electricalcurrent is applied across the thermoelectric device 94, thusestablishing an additional temperature difference between the hot faceand the cold face. By such provision and use of the thermoelectricdevice, the heat transfer is made more efficient. The fins 112 alsorender the heat transfer process more efficient.

FIG. 5 shows certain details of the water extractor 98 of thisembodiment. The water extractor 98 comprises vanes 120, a drain 122 andthe previously mentioned output 58. These elements are arranged suchthat the cooled output air 104 passes over the vanes which swirl the airallowing liquid water to be extracted from the air by a scraperarrangement. The moisture exits the water extractor 98 in the form ofwater droplets 124 via the drain 122. This leaves cooled and driedoutput air 106 which exits the water extractor 98 (and hence the firstheat exchanger 51) via the output 58.

The heat being scavenged may be transferred, via the third heatexchanger 53 and the second heat exchanger 52, to air being recirculatedaround the air circulation path 26 via the second heat exchanger 52, toheat the equipment bay 18. Appropriate routing and recirculation of theair may be achieved by using appropriate settings of the various valvesystems 81-85 as described earlier above.) Additionally, or instead, theheat being scavenged may be transferred via the third heat exchanger 53to other parts of the turboprop-powered MALE aircraft 1 to performheating, for example via one or more further fluid circulation pathsthat pass via the third heat exchanger 53, to locations such as near tothe leading edges of the wings 4 and/or engine support pylons 6, forexample to prevent ice forming. Additionally, or instead, the heat beingscavenged may be transferred directly to other parts of theturboprop-powered MALE aircraft 1 to perform heating, for example atleast part of the feed stage 40 of the pressurised water-glycol heattransfer loop 28 may be positioned to run in locations such as near tothe leading edges of the engine support pylons 6, for example to preventice forming. (This latter example of the feed stage 40 of thepressurised water-glycol heat transfer loop 28 being arranged to runnear to the leading edges of the engine support pylons 6, for example toprevent ice forming, is particularly advantageous in view of the factthat the heat is being obtained by scavenging heat from the engines,thus the distance of travel between where the fluid gains the scavengedheat to where it is used to give up this heat to e.g. prevent iceforming, is very short and hence this process can be very efficient.)

It will be appreciated that the above described arrangements may bedeployed in many advantageous ways and combinations according to theflight conditions the turboprop-powered MALE aircraft 1 will undergobefore and during flight. Various examples of these flight conditions,and associated embodiments of processes of performing thermal managementof the turboprop-powered MALE aircraft 1 during such flights, will nowbe described with reference to FIGS. 6 to 11. It will be appreciatedthat the air temperatures given in the below examples are merely nominaltemperatures to allow further understanding, and are not limiting to theinvention. Also, it will be appreciated that for clarity various examplemission profiles are described below in terms of simplified discretephases, although in practise such phases would typically merge into eachother and vary more subtly.

FIG. 6 is a schematic illustration of a first example of a missionprofile 150 that the turboprop-powered MALE aircraft 1 may undergo. Forease of reference, this example may be considered as being one in whichthe turboprop-powered MALE aircraft 1 experiences “hot” conditions.

During a first phase 151, the turboprop-powered MALE aircraft 1undergoes ground conditioning, i.e. the turboprop-powered MALE aircraft1 is on the ground and is connected to an air cooling trolley and, asdesired, other ground conditioning/processing equipment. In this examplethe air temperature on the ground is 50° C.

During a second phase 152, the air cooling trolley and any other groundconditioning equipment is disconnected, and the turboprop-powered MALEaircraft 1 carries out taxiing, taking-off and climbing. In this examplethe air temperature experienced by the turboprop-powered MALE aircraft 1during this phase is 50° C. whilst taxiing and taking off and thendecreases gradually to 10° C. as the turboprop-powered MALE aircraft 1climbs to its mission altitude.

During a third phase 153, the turboprop-powered MALE aircraft 1 flies atits mission altitude, where in this example the air temperature is 10°C.

During a fourth phase 154, the turboprop-powered MALE aircraft 1descends, lands, and taxis in, thereby experiencing in this example anair temperature that rises from 10° C. to 50° c. over the course of thefourth phase 154.

In this first mission profile 150, the duration of each of the firstphase 151 and third phase 153 is very long compared to the duration ofthe second phase 152 and the fourth phase 154. This will typically arisedue to the particular characteristics of the turboprop-powered MALEaircraft 1, in particular its long endurance flight capability. Forexample, the third phase 153 may be longer than 12 hours, whereas thesecond phase 152 may be less than 30 minutes.

FIG. 7 is a process flowchart showing certain steps performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft 1 during the above described firstexample of a mission profile 150.

At step s2, during the above described first phase 151 of the firstexample of a mission profile 150, the thermal control system 22 isoperated such as to use ground cooling air from the ground coolingtrolley input via the air inlet 24 to cool the first heat exchanger 51and the equipment bay 18, and after performing such cooling on a singlepass through the equipment bay 18, this air is allowed to exit thermalcontrol system 22 via the air outlet 25, by using appropriate settingsof the various valve systems 81-85 as described earlier above. Duringthis process the heat-storage wax in the heat sink layers 92 of thefirst heat exchanger 51 are solidified by virtue of being cooled by thereadily available cool air from the ground cooling trolley. Optionally,this process may be enhanced by applying electrical power to thethermo-electric device 94.

At step s4, during the above described second phase 152, air isrecirculated in the air circulation path 26 and is cyclically passed viathe first heat exchanger to cool the equipment bay 18, by usingappropriate settings of the various valve systems 81-85 as describedearlier above. During this process, for the relatively short time of theabove described second phase 152, due to the specific/latent heat of thesolidified heat-storage wax, as the heat-storage wax is melted by thewarm air, the first heat exchanger is able to efficiently cool therecirculating air that is being used to cool the equipment bay 18.

At step s6, during the above described third phase 153, the external airis now cool enough to be used for cooling the equipment bay 18, and soram air scooped in via the air inlet 24 is used to cool the first heatexchanger 51 and the equipment bay 18, and after performing such coolingon a single pass by, this air is allowed to exit thermal control system22 via the air outlet 25, by using appropriate settings of the variousvalve systems 81-85 as described earlier above. During this process theheat-storage wax in the heat sink layers 92 of the first heat exchanger51 are solidified by virtue of being cooled by the readily availablecool air from the ram air scoop. (Note, in flights where the first heatexchanger 51, and hence the heat storage properties of the heat storagewax, will not be needed for the following fourth phase 154, then duringthe third phase 153 the ram air may be used just to cool the equipmentbay 18, thereby requiring less ram air, or tolerating warmer ram air,for example.)

At step s10, during the above described fourth phase 154, air isrecirculated in the air circulation path 26 and is cyclically passed viathe first heat exchanger 51 to cool the equipment bay 18, by usingappropriate settings of the various valve systems 81-85 as describedearlier above. During this process, for the relatively short time of theabove described fourth phase 154, due to the specific/latent heat of thesolidified heat-storage wax, as the heat-storage wax is melted by thewarm air, the first heat exchanger 51 is able to efficiently cool therecirculating air that is being used to cool the equipment bay 18.

FIG. 8 is a schematic illustration of a second example of a missionprofile 160 that the turboprop-powered MALE aircraft 1 may undergo. Forease of reference, this example may be considered as being one in whichthe turboprop-powered MALE aircraft 1 experiences “mild, but cold ataltitude” conditions.

During a first phase 161, the turboprop-powered MALE aircraft 1 is onthe ground and undergoes general ground conditioning, but since in thisexample the air temperature on the ground is 20° C., no specialisedcooling or heating is required.

During a second phase 162, the turboprop-powered MALE aircraft 1 carriesout taxiing, taking-off and climbing. In this example the airtemperature experienced by the turboprop-powered MALE aircraft 1 duringthis phase is 20° C. whilst taxiing and taking off and then decreasesgradually to −10° C. as the turboprop-powered MALE aircraft 1 climbs toits mission altitude. In this example during this second phase 162 nospecialised cooling or heating is required (e.g. in this simplifiedaccount let us assume the climb is sufficiently rapid that any coolinginduced by the external air is insignificant until the following thirdphase is reached).

During a third phase 163, the turboprop-powered MALE aircraft 1 flies atits mission altitude, where in this example the air temperature is −10°C. In this example we assume this is the only phase where heating isrequired to be performed, which heating will be described in more detailbelow with reference to FIG. 9.

During a fourth phase 164, the turboprop-powered MALE aircraft 1descends and lands, thereby experiencing in this example an airtemperature that rises from −10° C. to 20° c. over the course of thefourth phase 164.

In this second mission profile 160, the duration of the third phase 163is very long compared to the duration of each of the other phases. Thiswill typically arise due to the particular characteristics of theturboprop-powered MALE aircraft 1, in particular its long enduranceflight capability/requirement. For example, the third phase 163 may belonger than 12 hours, even longer than 24 hours. More generally, it isdesirable in this type of aircraft for the available endurance to be aslong as possible.

FIG. 9 is a process flowchart showing a certain step performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft 1 during the above described secondexample of a mission profile 160.

In this thermal management process, the step s8 is performed during theabove described third phase 163 of the second example of a missionprofile 160. At step s8, heating is performed using heat that is beingscavenged from one or more hot parts of the gas turbine engine 8, forexample in the manner described earlier above employing the thermalblankets 44. The heat being scavenged may be transferred, via the thirdheat exchanger 53 and the second heat exchanger 52, to air beingrecirculated around the air circulation path 26 via the second heatexchanger 52, to heat the equipment bay 18. (Again, appropriate routingand recirculation of the air is achieved by using appropriate settingsof the various valve systems 81-85 as described earlier above.)Additionally, or instead, the heat being scavenged may be transferredvia the third heat exchanger 53 to other parts of the turboprop-poweredMALE aircraft 1 to perform heating, for example via one or more furtherfluid circulation paths that pass via the third heat exchanger 53, tolocations such as near to the leading edges of the wings 4 and/or enginesupport pylons 6, for example to prevent ice forming. Additionally, orinstead, the heat being scavenged may be transferred directly to otherparts of the turboprop-powered MALE aircraft 1 to perform heating, forexample at least part of the feed stage 40 of the pressurisedwater-glycol heat transfer loop 28 may be positioned to run in locationssuch as near to the leading edges of the engine support pylons 6, forexample to prevent ice forming. (This latter example of the feed stage40 of the pressurised water-glycol heat transfer loop 28 being arrangedto run near to the leading edges of the engine support pylons 6, forexample to prevent ice forming, is particularly advantageous in view ofthe fact that the heat is being obtained by scavenging heat from theengines, thus the distance of travel between where the fluid gains thescavenged heat to where it is used to give up this heat to e.g. preventice forming is very short and hence this process can be very efficient.)

FIG. 10 is a schematic illustration of a third example of a missionprofile 170 that the turboprop-powered MALE aircraft 1 may undergo. Forease of reference, this example may be considered as being one in whichthe turboprop-powered MALE aircraft 1 experiences “hot ground conditionsand a mixture of hot and cold conditions at altitude”.

During a first phase 171, the turboprop-powered MALE aircraft 1undergoes ground conditioning, i.e. the turboprop-powered MALE aircraft1 is on the ground and is connected to an air cooling trolley and asdesired, other ground conditioning/processing equipment. In this examplethe air temperature on the ground is 50° C.

During a second phase 172, the air cooling trolley and any other groundconditioning equipment is disconnected, and the turboprop-powered MALEaircraft 1 carries out taxiing, taking-off and climbing. In this examplethe air temperature experienced by the turboprop-powered MALE aircraft 1during this phase is 50° C. whilst taxiing and taking off and thendecreases gradually to 10° C. as the turboprop-powered MALE aircraft 1climbs to its mission altitude.

During a third phase 173, the turboprop-powered MALE aircraft 1 flies afirst part of its mission, during which in this example the airtemperature is 10° C.

During a fourth phase 174, the turboprop-powered MALE aircraft 1 flies asecond part of its mission, during which in this example the airtemperature is −10° C. Various reasons are possible for why there is adifferent external air temperature during the second part of the missioncompared to the first part of the mission. One possibility is the secondpart of the mission takes place at a higher altitude than the firstpart. Another possibility is that due to the long flight durationcharacteristics of the turboprop-powered MALE aircraft 1, the first partof the mission is e.g. during the warmer day and the second part isduring the e.g. colder night. In this example we assume this fourthphase 174 is the only phase where heating is required to be performed,which heating will be described in more detail below with reference toFIG. 11.

During a fifth phase 175, the turboprop-powered MALE aircraft 1 descendsand lands, thereby experiencing in this example an air temperature thatrises from −10° C. to 50° c. over the course of the fifth phase 175.

In this third mission profile 170, the duration of each of the firstphase 171, the third phase 173, and the fourth phase 174 is very longcompared to the duration of the second phase 172 and the fifth phase175. This will typically arise due to the particular characteristics ofthe turboprop-powered MALE aircraft 1, in particular its long enduranceflight capability. For example, the third phase 173 and the fourth phasemay each be longer than 6 or 12 hours, whereas the second phase 172 maybe less than 30 minutes.

FIG. 11 is a process flowchart showing certain steps performed in athermal management process that may advantageously be employed by theturboprop-powered MALE aircraft 1 during the above described thirdexample of a mission profile 170.

At step s2, during the above described first phase 171 of the thirdexample of a mission profile 170, the thermal control system 22 isoperated such as to use ground cooling air from the ground coolingtrolley input via the air inlet 24 to cool the first heat exchanger 51and the equipment bay 18, and after performing such cooling on a singlepass through the equipment bay 18, this air is allowed to exit thermalcontrol system 22 via the air outlet 25, by using appropriate settingsof the various valve systems 81-85 as described earlier above. Duringthis process the heat-storage wax in the heat sink layers 92 of thefirst heat exchanger 51 are solidified by virtue of being cooled by thereadily available cool air from the ground cooling trolley. Optionally,cooling of the heat storage wax may be enhanced by applying electricalpower to the thermo-electric device 94.

At step s4, during the above described second phase 172, air isrecirculated in the air circulation path 26 and is cyclically passed viathe first heat exchanger 51 to cool the equipment bay 18, by usingappropriate settings of the various valve systems 81-85 as describedearlier above. During this process, for the relatively short time of theabove described second phase 172, due to the specific/latent heat of thesolidified heat-storage wax, as the heat-storage wax is melted by thewarm air, the first heat exchanger is able to efficiently cool therecirculating air that is being used to cool the equipment bay 18.

At step s6, during the above described third phase 173, the external airis now cool enough to be used for cooling the equipment bay 18, and soram air scooped in via the air inlet 24 is used to cool the first heatexchanger 51 and the equipment bay 18, and after performing such coolingon a single pass through the equipment bay 18, this air is allowed toexit thermal control system 22 via the air outlet 25, by usingappropriate settings of the various valve systems 81-85 as describedearlier above. During this process the heat-storage wax in the heat sinklayers 92 of the first heat exchanger 51 are solidified by virtue ofbeing cooled by the readily available cool air from the ram air scoop.(Note, in flights where the heat-storage wax will not be needed for thelater fifth phase 155, then during the third phase 173 the ram air maybe used just to cool the equipment bay 18, thereby requiring less ramair, or tolerating warmer ram air, for example.)

At step s8, during the above described fourth phase 174, heating isperformed using heat that is being scavenged from one or more hot partsof the gas turbine engine 8, for example in the manner described earlierabove employing the thermal blankets 44. The heat being scavenged may betransferred via the third heat exchanger 53 and the second heatexchanger 52 to air being recirculated around the air circulation path26 via the second heat exchanger 52 to heat the equipment bay 18.(Again, appropriate routing and recirculation of the air is achieved byusing appropriate settings of the various valve systems 81-85 asdescribed earlier above.) Additionally, or instead, the heat beingscavenged may be transferred via the third heat exchanger 53 to otherparts of the turboprop-powered MALE aircraft 1 to perform heating, forexample via one or more further fluid circulation paths that pass viathe third heat exchanger 53 to locations such as near to the leadingedges of the wings 4 and/or engine support pylons 6, for example toprevent ice forming. Additionally, or instead, the heat being scavengedmay be transferred directly to other parts of the turboprop-powered MALEaircraft 1 to perform heating, for example at least part of the feedstage 40 of the pressurised water-glycol heat transfer loop 28 may bepositioned to run in locations such as near to the leading edges of theengine support pylons 6, for example to prevent ice forming. (Thislatter example of the feed stage 40 of the pressurised water-glycol heattransfer loop 28 being arranged to run near to the leading edges of theengine support pylons 6, for example to prevent ice forming, isparticularly advantageous in view of the fact that the heat is beingobtained by scavenging heat from the engines, thus the distance oftravel between where the fluid gains the scavenged heat to where it isused to give up this heat to e.g. prevent ice forming is very short andhence this process can be very efficient. A corresponding advantage isobtained in further embodiments where engines are mounted on the wings,and icing prevention of the leading edges of those wings is performed bydirect routing of heat scavenged from the engine mounted nearby on therespective wing.)

At step s10, during the above described fifth phase 175, air isrecirculated in the air circulation path 26 and is cyclically passed viathe first heat exchanger 51 to cool the equipment bay 18, by usingappropriate settings of the various valve systems 81-85 as describedearlier above. During this process, for the relatively short time of theabove described fifth phase 175, due to the specific/latent heat of thesolidified heat-storage wax, as the heat-storage wax is melted by thewarm air, the first heat exchanger 51 is able to efficiently cool therecirculating air that is being used to cool the equipment bay 18.

In the processes of FIGS. 9 and 11, in addition to some of the heat thatis being scavenged from the gas turbine engines 8 being used to heat airin the air circulation path 26 in order to heat the equipment bay 18,some of the heat that is being scavenged may instead be directly orindirectly (via the third heat exchanger 53) transferred to other partsof the turboprop-powered MALE aircraft 1 to perform heating, for exampleat near to the leading edges of the engine support pylons 6 and/or thewings 4. This is performed at different times during a flight to whencool air is being used to cool the equipment bay 18 (or when no suchcooling takes place on a flight, or at least when only minimalhotspot-avoiding air conditioning is taking place). However, in otherembodiments, at the same time that heat that is being scavenged isdirectly or indirectly (via the third heat exchanger 53) beingtransferred to other parts of the turboprop-powered MALE aircraft 1 toperform heating, for example at near to the leading edges of the enginesupport pylons 6 and/or the wings 4, cooling of the equipment bay 18 maybe carried out, this may occur when the outside air temperature is coldenough that various parts of the turboprop-powered MALE aircraft 1 suchas leading edges need heating to prevent ice formation, yet at the sametime the equipment in the equipment bay 18 produces levels of heat highenough for the equipment to still need to be cooled even though theoutside of the turboprop-powered MALE aircraft 1 requires heating. Interms of the flowchart of FIG. 11, this implementation may be consideredto be shown in FIG. 11, but noting that although steps such as s8 ands10 have, for convenience and ease of understanding, been depicted asdiscrete temporally-sequential steps, nevertheless such process stepsmay in fact be performed simultaneously or at least overlapping to someextent temporally.

The various cooling and heating operations described above in e.g. theprocesses of FIGS. 7, 9 and 11 may be selected as required for a giventurboprop-powered MALE aircraft 1 according to the particular details ofa given flight. In other words, the turboprop-powered MALE aircraft 1 isequipped with the various elements described above, and these areselected to be used as required. This advantageously provides a “onesystem suits all” solution that can be fitted as standard to pluralaircraft and which renders the aircraft capable of operating indiffering environments and for differing roles. Another possibility,however, is to select in advance just certain parts of the abovedescribed overall system, for aircraft that are only intended to operatein a more limited range of conditions or roles. For example only theelements necessary to provide one or more of the types of heatingreferred to in the process of FIG. 9 may be included, i.e. without anyof the elements used for cooling. Another possibility, for example, isthat only the elements necessary to provide the cooling process of FIG.7 may be provided, i.e. without any of the elements for the heatingprocesses described. Yet another possibility is that elements forcooling and heating are provided, but the heating does not include thoseelements required to provide heating of the equipment bay. Within allthe whole range of possibilities, circumstances/requirements exist whereone or more of the heat exchangers may be omitted, or where the secondheat exchanger 52 and the third heat exchanger 53 may be implemented asa common heat exchanger (either along with, or without, depending on theparticular implementation, the first heat exchanger 51).

In the above embodiments the turboprop-powered MALE aircraft 1 is anunmanned aircraft, i.e. an unmanned air vehicle (UAV), however this neednot be the case, and in other embodiments the aircraft may be manned.

In other embodiments, more than one equipment bay is cooled/heated. Inother embodiments, areas or parts of the aircraft other than anequipment bay as such may be air cooled/heated, either in addition to orinstead of the above described equipment bay.

In the above described embodiments the various components of the thermalcontrol system are located together physically as well as operably.However, this need not be the case, and in other embodiments any one ormore of the above described components may be distributed at variouslocations within the aircraft.

Many elements and details described with regard to the above particularembodiments, whilst individually or in combination advantageous, arenevertheless not essential. For example, in simplified embodiments, anyone or more of the following detailed elements may be omitted.

For ground cooling, any appropriate ground-based cooling apparatus maybe used, not just a ground cooling trolley as such.

Separate air inlets may be provided for ground cooling and ram air scoopcooling.

The air outlet 25 may be omitted, with air that has extracted heat fromthe equipment bay being allowed to escape directly from the equipmentbay by leakage or other general diffusion or dissipation ways withoutreturning to the thermal control system 22. If required, when the air isto be recirculated, active methods (e.g. a fan or a suction method) maybe employed to direct the air from the equipment bay 18 back into thethermal control system 22.

The particular valve systems arrangement shown may be replaced by anyappropriate routing and control process. For example, a simpler routingarrangement and/or less valve systems may be employed, for example iffewer options are required to be available in a given turboprop-poweredMALE aircraft. Another possibility is that the various functionsprovided by different parts of the single air circulation path 26 (withthe different functions selected by different settings of the valvesystems) can instead be implemented by providing plural separatealternative air circulation paths, e.g. each with their own entry intoand/or out of the equipment bay, and operationally selecting betweenthem as required.

The choice of the heat transfer liquid in the pressurised water-glycolheat transfer loop 28 is not essential, and instead any otherappropriate heat transfer liquid may be used. Possibilities include PAO(poly alpha olefin), and “Coolanol” (trademark).

The thermal blankets 44 need not necessarily be placed on the enginecasings 14 and/or the exhaust ducts 16, and may instead be placedanywhere heat can be scavenged, e.g. at other locations in or near thegas turbine engines.

One particular advantage of the use of the thermal blankets 44 is thatthey can be applied in a retrofit fashion to existing engine designs.However, the heat scavenging apparatus need not be the above describedthermal blankets, and instead (or in addition) other suitable types ofheat scavenging/heat extraction equipment may be used, such as engineoil heat exchangers.

Another possibility is that heat may instead or in addition be scavengedfrom the heat given out by the engines by scavenging the heat from theexhaust gases of the engines. For example, exhaust gas to water-glycolheat exchangers may be incorporated into modified engine exhaust ducts.

In further embodiments, heat provided when required to the aircirculation path 26 for heating the equipment bay 18 is obtained bybeing extracted in conventional fashion from the propulsion power, e.g.by taking bleed air form the gas turbine engine or by extracting shaftpower. Although in these embodiments the advantages related to avoidingloss of propulsion power are no longer achieved, nevertheless theoverall arrangement is still advantageous compared to prior artarrangements, since a simple combined air heating/air coolingarrangement for the equipment bay is provided, that provides not onlythe advantages of the above described cooling aspects, but alsoadditionally makes use of the air flow routes etc. for the equipment baywhen the equipment bay becomes too cold. Thereby a flexible thermalmanagement solution for a turboprop-powered MALE aircraft is providedthat can nevertheless accommodate widely varying environmentalcircumstances and different mission profiles.

In some of the embodiments described above, scavenged heat is used towarm parts of the aircraft, for example leading edges of wings and/orpylons to prevent icing. In further embodiments, such heating for theprevention of icing is in addition or instead applied to other areas ofthe aircraft, for example to engine air intake lips.

More generally, in yet further embodiments, the scavenged heat is usedin addition or instead to heat the fuel being carried by the aircraft,e.g. by delivering some or all of the scavenged heat to, or to thevicinity of, one or more fuel tanks or fuel supply lines of theaircraft. This will improve the temperature control of the fuel, ande.g. prevent or diminish fuel waxing.

In the above described embodiments, the configuration of the aircraft,and in particular the configuration of the engine positions, is asdescribed above and shown with reference to FIG. 1, i.e. one engine isprovided on each of two laterally extending engine-mounting pylons.However, in other embodiments, any other configuration may be used, asdesired. For example, one possibility is that a configuration of similarlayout to the aircraft of FIG. 1 may be employed, but with more than oneengine on each pylon. For example, another possibility is that, insteadof or in addition to engines mounted on the pylons, an engine may bemounted on a tail fin. Another possibility is that, in addition to orinstead of engines mounted on engine mounting pylons and/or a tail fin,one or more engines may be mounted on each wing.

The fins 112 of the thermoelectric devices 94 may be omitted. Indeed,the thermoelectric devices 94 themselves may be omitted.

Instead of plural air flow channels 90 and plural heat sink layers 92,there may be just one of either or both of these items. Also, it will beappreciated that the particular number of channels and heat sinks shownin Figure in FIG. 3 is merely one possible example that may be used.

The air flow channels 90 of the first heat exchanger 51 need not containopen-cell aluminium foam. One possibility is for the air flow channels90 to contain a different type of thermally conducting three-dimensionalmesh arrangement that allows air to flow. Another possibility is thatthe air flow channel can simply be hollow.

The heat sink layers 92 of the first heat exchanger 51 need not compriseopen-cell aluminium foam. One possibility is the heat-storage wax (orother heat-storage material) may be combined with a different type ofthree-dimensional thermally conducting mesh arrangement. Anotherpossibility is the heat sink layer 92 may comprise just heat-storage wax(or other heat-storage material).

Any suitable heat-storage wax may be used, for example paraffin wax.Another possibility is that instead of heat-storage wax, other suitabletypes of heat-storage material may be used.

Other items, such as the insulation jacket 96, the water extractor 98,and so on, are also of course not essential, and may be omitted. Whenincluded, the details of, or the types of, such items may be differentto those described in the above embodiments. For example, differenttypes of water extractor may be employed compared to the one describedabove.

Similarly, many details, such as the shape of the first heat exchanger51, the air flow channels 90 and the heat sink layers 92 etc. may bedifferent. For example, a circular cross-sectional design may beimplemented instead of the planar design shown in FIGS. 3 and 4.

It will be appreciated that the various embodiments described hereintend to provide, inter alia, one or more of the following advantages.

At least some of the embodiments provide a means for capturing heat thatreduces or removes completely any need to extract heat from the powerbeing provided for propulsion. This is advantageous in view of the longendurance requirement for a turboprop-powered MALE aircraft, i.e.propulsion power can be kept for contributing to extended flightendurance.

One particular advantage of the use of the thermal blankets 44 is thatthey can be applied in a retrofit fashion to existing engine designs.

At least some of the embodiments provide a cooling arrangement that doesnot need to use active air conditioning.

At least some of the embodiments provide an intermediate cooling processfor an intermediate part of the mission profile, which intermediate partoccurs between parts that are relatively straightforward to service interms of providing conventional cooling arrangements. In theintermediate part, the aircraft may be required to be stationary,perform taxiing to a take-off position, perform take-off and then climbto an operational altitude, or by completing the final descent, approachto landing and taxi in. Advantageously some or all of the embodimentsprovide an efficient solution for the cooling operation during such anintermediate stage that does not unduly complicate the simple coolingprocesses that may be used for the other stages. Such solutions areparticularly efficient during their limited time of deployment, and thatparticular efficiency tends to be derived at least in part by making useof the very fact that it is only required to carry out the cooling rolefor a short limited duration of the intermediate stage, and moreoveradvantageously tends to be easily “reset” during e.g. the first stage ofthe turboprop-powered MALE aircraft's next mission or during cruise ataltitude with respect to preparing for the landing phase of thatmission.

At least some of the embodiments provide a thermal managementarrangement for a turboprop-powered MALE aircraft that is able toprovide several of the above solutions in a turboprop-powered MALEaircraft such that they can be selected individually or in combinationaccording to particular flight conditions a turboprop-powered MALEaircraft undergoes on the same or different flights.

In various of the above described embodiments, advantageousimplementations that additionally include cooling aspects including useof a particular type of air cooling heat exchanger, and/or use ofscavenged heat to provide heating via parts that are at other times usedfor cooling, are included. However, it is to be noted that in simplerembodiments, no cooling aspects, e.g. use of a particular type ofcooling heat exchanger, are included, and/or no use of scavenged heat toprovide heating via parts that are at other times used for cooling isincluded. In other words, it is to be appreciated that simpleembodiments of a turboprop-powered MALE aircraft comprise a thermalmanagement system that need perform only heat scavenging of heat fromone or more gas turbine engines of the aircraft and use of the scavengedheat to provide heating to the aircraft, for example for icingprevention and/or fuel heating.

1. A turboprop-powered medium altitude long endurance aircraft engine,comprising: a gas turbine engine configured for providing propulsionpower to an aircraft propeller; a heat scavenging apparatus configuredfor scavenging heat from the gas turbine engine; and a heating apparatusconfigured to provide scavenged heat for heating an aircraft.
 2. Aturboprop-powered medium altitude long endurance aircraft engineaccording to claim 1, wherein the heat scavenging apparatus comprises: athermal blanket.
 3. A turboprop-powered medium altitude long enduranceaircraft engine according to claim 2, wherein the thermal blanketcomprises: capillary tubing.
 4. A turboprop-powered medium altitude longendurance aircraft engine according to claim 1, comprising: an enginecasing for the engine, wherein the heat scavenging apparatus is placedon the engine casing of the engine.
 5. A turboprop-powered mediumaltitude long endurance aircraft engine according to claim 1,comprising: an engine air duct of the engine, wherein the heatscavenging apparatus is placed on or in the exhaust duct of the engine.6. A turboprop-powered medium altitude long endurance aircraft engineaccording to claim 1, wherein the heating apparatus comprises: acirculation path for routing heat directly to an aircraft location whereheating is to be performed.
 7. A turboprop-powered medium altitude longendurance aircraft engine according to claim 6, in combination with anaircraft, wherein the location in the aircraft where heating is to beperformed is a leading edge of an engine support pylon or a leading edgeof an engine-carrying wing.
 8. A turboprop-powered medium altitude longendurance aircraft engine according to claim 1, wherein the heatingapparatus comprises: a heat exchanger.
 9. A turboprop-powered mediumaltitude long endurance aircraft engine according to claim 8, whereinthe heat exchanger is configured to apply scavenged heat to heat air ina recirculating air path.
 10. A turboprop-powered medium altitude longendurance aircraft engine according to claim 9, in combination with anaircraft, wherein the recirculating air path is arranged to providecooling and/or heating to the aircraft.
 11. A turboprop-powered mediumaltitude long endurance aircraft engine according to claim 10,comprising: a further heat exchanger; and wherein: the further heatexchanger comprises a heat-storage material; the further heat exchangeris arranged to cool air recirculating around the air recirculation path;and the air recirculation path is arranged to allow the recirculatingair to provide aircraft cooling.
 12. A turboprop-powered medium altitudelong endurance aircraft engine according to claim 11, wherein thefurther heat exchanger is configured to allow the heat-storage materialto be cooled by a ground-based cooling apparatus during aircraftgrounding.
 13. A turboprop-powered medium altitude long enduranceaircraft engine according to claim 12, wherein the heat-storage materialhas a melting point selected so as to be rendered solid when cooling isto be performed by the ground-based cooling apparatus and/or by ram air.14. A turboprop-powered medium altitude long endurance aircraft engineaccording to claim 11, wherein the further heat exchanger is configuredsuch that the heat-storage material will be cooled by ram air duringaircraft flight.
 15. A turboprop-powered medium altitude long enduranceaircraft engine according to claim 14, wherein the heat-storage materialhas a melting point selected so as to be rendered solid when cooling isto be performed by a ground-based cooling apparatus and/or by the ramair.
 16. A turboprop-powered medium altitude long endurance aircraftengine according to claim 1, wherein the heating apparatus comprises: apressurised heat transfer liquid loop.
 17. A turboprop-powered mediumaltitude long endurance aircraft engine according to claim 16, incombination with an aircraft, wherein the pressurised heat transferliquid loop is configured to transfer scavenged heat to aircraft partsto perform icing prevention at those parts of the aircraft.
 18. A methodof thermally managing a turboprop-powered medium altitude long enduranceaircraft, the method comprising: scavenging heat from an aircraftengine; and performing heating of the turboprop-powered medium altitudelong endurance aircraft with heat scavenged from the engine of theaircraft.
 19. A method according to claim 18, comprising: cooling, withground-based cooling air, a heat exchanger, the heat exchanger having aheat-storage material; thereafter, passing recirculated air via the heatexchanger to cool the turboprop-powered medium altitude long enduranceaircraft; and thereafter cooling, with ram air, the heat exchanger.